Integrated air and vapor cycle cooling system

ABSTRACT

A cooling system in which an ACS (air cycle system) turbine may be driven by high pressure air from a turbo-fan engine and a VCS (vapor cycle system) having an evaporator and a VCS refrigerant compressor may be driven by the ACS turbine. Fluid of the chilled fluid reservoir, which may be chilled fuel, may be circulated through and cooled in the evaporator. In some embodiments, the ACS turbine may be coupled to the VCS refrigerant compressor by a magnetic coupling.

RELATED APPLICATIONS

The present application claims benefit of U.S. Provisional ApplicationNo. 62/212,201 filed Aug. 31, 2015.

BACKGROUND OF THE INVENTION

The present invention generally relates to cooling systems, and moreparticularly to systems that include both air and vapor cycle coolingsystems.

Future military aircraft will require greatly advanced capabilities inorder to ensure air superiority, survival in heavily defended airspace,and success against a wide range of potential targets. Thesecapabilities, including electronic attack and directed energy weapons,are expected to require as much as 10 times higher electric power levelsthan existing weapon systems.

While these power demands are significant, thermal management of theseadvanced systems will present an even greater challenge, due to lowcomponent efficiencies and waste heat quality. Low-observabilityrequirements, such as limitations on ram air availability, compositeaircraft skins which inhibit heat transfer, and higher efficiencyengines with less fuel flow available for cooling, further compound thischallenge.

Air cycle systems (ACS) offer the potential advantage of high pressureratios and lift temperatures, and thus can be used with hotter heatsinks than comparable vapor cycle systems (VCS). However, ACS are muchless efficient than VCS, resulting in higher power requirements andhigher heat rejection demands.

On the other hand, VCS systems are more efficient, but require lowerheat sink temperatures that are often not available without a dedicatedram air source, which still may be too hot to use during high speedflight.

Further, for the large thermal loads anticipated, the cooling systemitself will require substantial power to drive it. The use of electricpower would further tax the vehicle's electric power generationcapability and the power conditioning, motor controllers, and motorsneeded would be relatively heavy, expensive, and unreliable. A directmechanical drive may also be problematic as engines are not currentlyequipped for such high power take-off shaft loads, and incorporatingsuch into new engine designs involves a number of design compromisesincluding added weight, cost and mechanical complexity.

As can be seen, there is a need for an efficient cooling system that canbe used in aircraft environments of the type described.

SUMMARY OF THE INVENTION

In one aspect of the invention, a cooling system comprises aturbo-compressor that includes a turbine that drives a hermeticallysealed, two-stage compressor; wherein the turbine receives air from anengine; a condenser downstream of the turbo-compressor; wherein thecondenser receives a discharge flow from the two-stage compressor andfrom the turbine; an evaporator downstream of the condenser; and whereinthe evaporator is upstream of the two-stage compressor.

In another aspect of the invention, a cooling system comprises an aircycle system (ACS) turbine driven by high pressure air from a turbo-fanengine; a vapor cycle system (VCS) comprising an evaporator and a VCScompressor driven by the ACS turbine; a condenser cooled by dischargefrom the turbine; and a chilled fluid reservoir, wherein fluid of thechilled fluid reservoir is circulated through the evaporator.

In a further aspect of the invention, a cooling system comprises an aircycle system (ACS) turbine driven by air from a turbo-fan engine; avapor cycle system (VCS) comprising a VCS refrigerant compressor drivenby the ACS turbine; a condenser cooled by ACS turbine discharge; and anevaporator; wherein the ACS turbine is magnetically coupled to the VCSrefrigerant compressor.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdrawings, description, and claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a simplified system schematic showing an example embodiment ofthe cooling system of the present invention.

FIG. 2 is a simplified system schematic showing another exampleembodiment of the cooling system of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

The following detailed description is of the best currently contemplatedmodes of carrying out exemplary embodiments of the invention. Thedescription is not to be taken in a limiting sense, but is made merelyfor the purpose of illustrating the general principles of the invention,since the scope of the invention is best defined by the appended claims.

Various inventive features are described below that can each be usedindependently of one another or in combination with other features.However, any single inventive feature may not address any of theproblems discussed above or may only address one of the problemsdiscussed above. Further, one or more of the problems discussed abovemay not be fully addressed by any of the features described below.

Generally, the present invention provides a hybrid air cycle system witha vapor cycle system by combining an ACS turbine and VCS compressor aspart of a single turbo-compressor. A “simple-cycle” ACS, with cooledengine fan air expanded across the turbo-compressor turbine providesboth shaft power to drive a vapor cycle compressor and a cool turbinedischarge to use as the vapor cycle heat sink. The VSC includes thecompressor driven by the turbine, which compresses a refrigerant gas; acondenser which rejects heat from the refrigerant gas to the turbinedischarge thereby condensing the refrigerant gas; an expansion valvewhich reduces the pressure of a liquid refrigerant from the condenser;and an evaporator which absorbs heat from the thermal loads into theliquid refrigerant causing it to evaporate.

In the case of a variable-cycle engine, the present invention providesengine fan air that can be cooled by the engine's third stream,providing a high-pressure cool air source to drive the turbo-compressor.On conventional turbofan engines the fan air can be used to cool theengine bleed air.

However, at high fan air temperatures, a greater pressure ratio acrossthe turbine can be required to bring the turbine discharge down to a lowenough temperature to condense the VCS refrigerant. Alternately,aircraft fuel or a separate ram air source can be used as needed tofurther cool the bleed air.

The terms “directly” or “direct”, in the context of two components ofthe system herein, is intended to mean that there is an absence of athird system component between the two components other than ducting toenable flow of a refrigerant or the like.

With reference now to FIG. 1, embodiments of the present inventionbroadly provide a cooling system 10 that may include an air cycle system42 and a vapor cycle system 40. In one embodiment, the cooling system 10may be used with a variable-cycle engine 12 that may be a turbo-fan typejet aircraft engine having an inlet fan 14 and a compressor 16, whichare connected by concentric shafts 18 to low- and high-pressure turbinesections 22. The compressor 16 and turbine sections 22 may befluidically coupled by a combustion chamber 20.

Engine fan air 24 from the variable-cycle engine 12 may be directedthrough a heat exchanger 26, which may be a fan air cooler, or the like,where it may be cooled by a 3rd stream air flow 28. The 3rd stream airflow available in a variable-cycle engine 12 is well-known in the art asa means for contributing cooling capability for aircraft systems andengine components. The resulting high-pressure air 30, which may be nowcooled to a moderate temperature, may be metered through a flow controlvalve 32 and into a turbine 34 of an integrated turbo-compressor 41, theturbine being integrated into the turbo-compressor configuration asdescribed more fully below. The flow control valve 32 may be used toadjust the speed of the turbine 34 of the turbo-compressor, as needed,to accommodate variations in thermal loads, for instance, of aircraft,or other environment, in which the cooling system 10 may be installed. Acool discharge air stream 38 of the turbine 34 can provide a cool airstream 38 which may be used as a heat sink for the VCS 40, as describedin greater detail below. The combination of engine fan air 24 cooled by3rd stream air flow 28 and expanded across the turbine 34 to produce acool discharge air stream 38 forms what is commonly referred to as a“simple-cycle” ACS 42.

The temperature at the outlet of the turbine 34 can be cool enough tocondense the refrigerant in a condenser 52 under all operatingconditions. If 3rd stream air 28 is not available, or if it is not coolenough to provide sufficient cooling, an additional heat sink, forexample, ram air, or aircraft fuel, may be utilized. Alternately, oradditionally, a higher pressure ratio across the turbine 34, such asmight be achieved with a multi-stage turbine 34′, as shown in theembodiment of FIG. 2, may be employed.

The turbine 34 of the turbo-compressor can drive the VCS 40 through ashaft with a seal or, as in the embodiment illustrated, the turbine 34can drive the VCS 40 through a magnetic coupling 44. The magneticcoupling 44 allows the shaft of a VCS refrigerant compressor 48 of theVCS 40 to be hermetically sealed to limit refrigerant leakage andeliminate the need for periodic refrigerant servicing. The magneticcoupling 44 can be replaced with an ordinary shaft seal if the resultingdecrease in refrigerant service intervals may be acceptable in theparticular environment in which the cooling system 10 may be employed.

The VCS 40 may further include a condenser 52, downstream of the turbine34, to receive a refrigerant gas 50 from the VCS refrigerant compressor48 to condense it to a high-pressure liquid refrigerant 56. An expansionvalve 58, downstream of the condenser 52, can reduce the pressure of therefrigerant 56. An evaporator 60, downstream of the expansion valve 58,can exchange heat with the fluid/fuel from a chilled fuel tank 64 and apump 66, which fluid/fuel, after heat exchange, can return to the fueltank 64. From the evaporator 60, a gas refrigerant 63 can flow to theVCS refrigerant compressor 48.

More particularly, the VCS refrigerant compressor 48 may compress arefrigerant gas 63 and supply that compressed refrigerate gas 50 to thecondenser 52. The condenser 52 can remove heat from the refrigerant gas50 and expel the heat to the turbine discharge 38 generating an exhaustgas 54. The condenser 52 may be cooled by turbine discharge 38, asshown, or other air flows; however, generally, in aircraft embodiments,3rd stream air may be too hot and ram air scoops increase drag and radarcross-section and may not be available during ground operation. Theexhaust gas 54 from the condenser 52 can be fed back into the 3rd streamair 28, if desired (not shown).

Thermal loads 62 may be carried to the evaporator 60 in liquid coolantsor gases such as air. However, in the embodiment of FIG. 1, theevaporator 60 can also be used to cool aircraft fuel which can then bestored in a dedicated chilled fuel tank 64. This chilled fuel can thenbe used to cool the thermal loads 62, other aircraft systems, or thelike. The fuel from the chilled fuel tank 64 can be pumped by a fuelpump 66 through the evaporator 60, where it may be cooled and returnedto the chilled fuel tank 64.

If and when needed, the fuel in the chilled fuel tank 64 can be directedto the main engine fuel boost pump (not shown) by opening fuel controlvalve 68. This approach offers the advantage that a relatively largeamount of thermal energy can be absorbed and stored. This thermalstorage capacity allows the integrated air and vapor cycle coolingsystem to be sized for average thermal loads over some period of timerather than for maximum peak loads, since the average thermal load mayoften be much lower than the peak loads, particularly in cases where theheat may be generated by directed energy weapons which do not operatecontinuously.

Unlike wax-based thermal storage systems or the other cooling fluids,using fuel for thermal storage offers the advantage of also being ableto use the fuel for propulsion. Of course, once the chilled fuel isburned in the engine it is no longer available for thermal storage, butas long as all the other fuel on the aircraft is used first, the thermalstorage capability can be maintained until near the end of the flight,such as while returning to base after a mission when thermal loads havemoderated.

Using chilled fuel provides potentially massive energy storage forextended operation with high thermal loads, or for temporaryinterruptions in cooling system operation, such as when maximum engineperformance may be required (e.g. during take-offs), or in the event ofa cooling system failure. In addition, chilled fuel supports steadyevaporator temperatures for simplified VCS control, while avoiding addedweight and volume of a dedicated thermal storage system. As mentioned,it also can be used for propulsion during egress when thermal loads arereduced.

With reference additionally now to FIG. 2, another example embodiment ofthe present invention is illustrated, showing at least some possiblealternative structures and devices for making and using a cooling system10′, which can be employed in aircraft or other environments. Thecooling system 10′ in this embodiment may be used in conjunction with aconventional fan-jet engine 12′. The conventional fan-jet engine 12′ hasan input fan 14′ and a compressor 16′, which are connected by concentricshafts 18′ to a low- and high-pressure turbine sections 22′. Thecompressor 16′ and turbine 22′ are fluidically coupled by a combustionchamber 20′.

Bleed air 25′ from the conventional fan-jet engine 12′ may be directedthrough a bleed air cooler 27′ where it may be cooled by engine fan air24′. The resulting high-pressure air 30′, which may now cooled to amoderate temperature, may be metered through a flow control valve 32′and into a two-stage turbine 34′, which may be integrated into aturbo-compressor configuration as described more fully below. Thetwo-stage turbine 34′ configuration may be used, for example, to achievesufficiently cool temperatures in the condenser 52′ or to reduce itssize.

The flow control valve 32′ may be used to adjust the speed of thetwo-stage turbine 34′, as needed, to accommodate variations in thermalloads, for instance, of aircraft, or other environment, in which thecooling system 10′ may be installed. The cool discharge air stream 38′of the two-stage turbine 34′ provides a cool air stream which may beused as a heat sink for the VCS 40′, as described in greater detailbelow. The combination of bleed air 25′ cooled by the engine fan air 24′and expanded across the two-stage turbine 34′ to produce a cooldischarge air stream 38 forms a simple-cycle ACS 42′.

The two-stage turbine 34′ drives the VCS 40′ through a magnetic coupling44′. In the embodiment of FIG. 2, a two-stage VCS refrigerant compressor48′ may be employed in the VCS 40′. The magnetic coupling 44′ allows theshaft of the two-stage VCS refrigerant compressor 48′ of the VCS 40′ tobe hermetically sealed to limit refrigerant leakage and eliminate theneed for periodic refrigerant servicing.

The two-stage VCS refrigerant compressor 48′ of the VCS 40′ compresses arefrigerant gas 50′ and directs it to a condenser 52′. The condenser 52′removes heat from the refrigerant gas 50′ and expels the heat to theturbine discharge air stream, generating the exhaust gas 54′. Thecondenser 52′ may be cooled by the cool turbine discharge air stream38′, as shown. The condenser 52′ condenses the refrigerant gas 50′ intoa high-pressure liquid refrigerant 56′. The pressure of the refrigerant56′ may be then reduced across an adjustable expansion valve 58′ anddirected into an evaporator 60′ where it can absorb heat from thermalloads 62′, for example, from an aircraft (not shown) in which the systemmay be installed. The output from the evaporator 60′ may be returned tothe two-stage VCS refrigerant compressor 48′ in liquid/gas form 63′ tocomplete the cycle. This combination of two-stage VCS refrigerantcompressor 48′, condenser 52′, adjustable expansion valve 58′, andevaporator 60′ comprise the VCS 40′.

In the embodiment of FIG. 2, the thermal loads 62′ are cooled in a heatexchanger 70′ through which coolant from a coolant reservoir 72′ may becirculated. The coolant may be pumped through the evaporator 60′,coolant reservoir 72′, and heat exchanger 70′ by a coolant pump 74′, asshown. Suitable liquid coolants might include polyalphaolephin (PAO),water, or water mixtures such as ethylene glycol and water (EGW) orpropylene glycol and water (PGW). These liquids could be used, forexample, to cool high-power aircraft electronics including directedenergy weapons, engine components, and actuators. Gases such as airwould might be air used to cool avionics and for cockpit cooling andpressurization.

The various design option embodiments described above are not intendedto be exhaustive or complete as other design options will be apparent tothose skilled in the art. For example, the air stream provided to theturbine 34 of the turbo-compressor in FIG. 1 or two-stage turbine 34′ inFIG. 2 may be derived from fan air, engine bleed air, 3^(rd) stream air,or another air source. The cooling medium used in the heat exchanger 26in FIG. 1 or the bleed air cooler 27′ in FIG. 2 may be provided by fanair, 3^(rd) stream air, ram air, fuel, or another cooling source.

It should be understood, of course, that the foregoing relates toexemplary embodiments of the invention and that modifications may bemade without departing from the spirit and scope of the invention as setforth in the following claims.

We claim:
 1. A cooling system, comprising: an air cycle system (ACS)comprising a two-stage turbine driven by air from a turbo-fan engine; avapor cycle system (VCS) comprising an evaporator, a two-stagecompressor driven by the turbine, and a condenser cooled by dischargefrom the turbine; wherein the evaporator receives a liquid refrigerantfrom the condenser; wherein the condenser receives a gas refrigerantfrom a second stage of the compressor; wherein a first stage of thecompressor receives a liquid/gas refrigerant from the evaporator; aliquid coolant reservoir directly downstream of the evaporator; a heatexchanger directly downstream of the liquid coolant reservoir; whereinthe heat exchanger is upstream of the evaporator with only a coolantpump and ducting between the heat exchanger and the evaporator; whereinthe heat exchanger receives a thermal load from outside of the coolingsystem; and wherein liquid coolant from the from the heat exchanger isre-circulated back through the evaporator and then through the liquidcoolant reservoir.
 2. The cooling system of claim 1 further comprising:an expansion valve between the condenser and the evaporator.
 3. Thecooling system of claim 1 further comprising a magnetic coupling betweenthe turbine and the compressor.
 4. The cooling system of claim 1 whereinthe liquid coolant is engine fuel.
 5. The cooling system of claim 1wherein the liquid coolant is selected from the group consisting ofpolyalphaolephin (PAO), water, or water mixtures such as ethylene glycoland water (EGW) or propylene glycol and water (PGW).
 6. The coolingsystem of claim 1 wherein the thermal load is selected from the groupconsisting of air for cockpit cooling, aircraft electronics loads,aircraft avionics, and directed energy weapons.
 7. The cooling system ofclaim 1, wherein the air from the turbo-fan engine is provided by athird stream air.
 8. The cooling system of claim 1 further comprising acontrol valve between the turbo-fan engine and the turbine to control anair flow to the turbine for adjusting a speed of the turbine as neededaccording to variations in aircraft thermal loads.
 9. A cooling system,comprising: an air cycle system (ACS) comprising a two-stage turbinedriven by air from an engine; wherein a first stage of the turbinereceives the engine air; a vapor cycle system (VCS) comprising atwo-stage compressor driven by the turbine, a condenser cooled by asecond stage turbine discharge, and an evaporator; wherein the two-stageturbine drives a second stage of the compressor and also drives, via thesecond stage of the compressor, a first stage of the compressor; whereinthe first stage of the compressor receives a discharge from theevaporator; a magnetic coupling directly between the first stage of theturbine and the second stage of the compressor; wherein there is anabsence of system components between the first stage of the turbine andthe second stage of the compressor other than a shaft with a magneticcoupling; and wherein the condenser receives a discharge flow directlyfrom the second stage of the compressor and receives a discharge flowdirectly from a cooling stage of the turbine.
 10. The cooling system ofclaim 9 wherein the magnetic coupling hermetically isolates therefrigerant compressor.
 11. The cooling system of claim 9 wherein theengine is a variable cycle engine.
 12. The cooling system of claim 9wherein the engine is a turbo-fan engine.
 13. The cooling system ofclaim 9 further comprising a control valve between the turbo-fan engineand the turbine to control an air flow to the turbine for adjusting aspeed of the turbine according to variations in aircraft thermal loads.